Reaction type rocket propulsion systems are often used to propel vehicles by burning a propellant in a combustion chamber and allowing the products of combustion to escape through a nozzle. According to the law of conservation of momentum, the change in momentum, or impulse, of the escaping products must produce an equal change in the momentum of the vehicle from which the products escape. Consequently, the reaction force generated by the burning propellant imparts thrust upon the vehicle.
Reference is now made to FIG. 1 showing a schematic representation of a typical rocket propulsion system 10 of the PRIOR ART. The prior art rocket propulsion system 10 includes a combustion chamber 12 and a nozzle 18. When propellant 14 contained within the combustion chamber is ignited, the products 16 of combustion are ejected from the nozzle 18.
The impulse that can be generated from a conventional rocket thruster such as shown in FIG. 1, depends upon the quantity and quality of propellant contained within the combustion chamber and the geometry of the nozzle. For a given amount of propellant, the thrust of a rocket motor can be written as F=CFPcAt where Pc, At are the pressure in the combustion chamber and nozzle throat cross-section area, respectively. The thrust coefficient CF is typically about 1.6 and can reach 1.8 in space. In order to achieve high pressure in the combustion chamber for a given flow rate, the nozzle throat cross-section area in a typical rocket motor is relatively small.
The fact that the combustion gases leave the vehicle at high speed and temperature indicates that a significant portion of the energy of the propellant is lost.
The need remains therefore for an efficient rocket propulsion system which is able to convert a greater proportion of the energy of the propellant into kinetic energy of vehicle. Embodiments described herein address this need.